A turbine shroud segment for a gas turbine engine having an annular gas path extending about an engine axis. The engine has a turbine rotor mounted for rotation about the axis and having a plurality of blades extending into the gas path. The turbine shroud segment includes a body extending axially between a leading edge and a trailing edge and circumferentially between a first and second lateral edges. The body has a radially outer surface and a radially inner surface. The radially outer surface includes a textured surface exposed to a cooling flow. The radially inner surface defines an outer flow boundary surface of the gas path next to a tip of one of the blades. A cooling flow passageway is defined in the body and extends axially between one or more cooling inlets receiving the cooling flow from the textured surface and one or more cooling outlets.

An embodiment of a torch igniter for a combustor of a gas turbine engine includes a combustion chamber oriented about an axis, a cap defining the axially upstream end of the combustion chamber and situated on the axis, a tip defining the axially downstream end of the combustion chamber, an igniter wall extending from the cap to the tip and defining a radial extent of the combustion chamber, a structural wall coaxial with and surrounding the igniter wall, an outlet passage defined by the igniter wall within the tip, wherein the outlet passage fluidly connects the combustion chamber to the combustor of the gas turbine engine, and a cooling system. The cooling system has an air inlet, a cooling channel, and an aperture. The cooling channel forms a flow path having a first axial section, a second axial section, a radially inward section, and a radially outward section.

An engine component assembly includes a first engine component having a hot surface in thermal communication with a hot combustion gas flow and a cooling surface with at least one cavity. A second engine component is spaced from the cooling surface, and includes at least one cooling aperture. The cooling aperture is arranged such that cooling fluid impinges on the cooling surface at an angle.

The invention relates to an air intake of an aircraft turbojet engine nacelle, extending along an axis X, in which an air flow circulates from upstream to downstream, the air intake extending circumferentially around the axis X and comprising an inner wall, which faces the axis X in order to guide an inner air flow, and an outer wall, which is opposite the inner wall, for guiding an external air flow, the walls being connected by a leading edge and an inner partition so as to delimit an annular cavity. The air intake comprises means for injecting at least one hot air flow into the inner cavity and at least one ventilation orifice formed in the outer wall in order to allow the hot air flow to escape after heating the internal cavity, the air intake comprising at least one disruption member of the external air flow, positioned upstream of the ventilation orifice, which extends outwardly from the outer wall.

A turbine shroud segment for a gas turbine engine having an annular gas path extending about an engine axis. The engine has a turbine rotor mounted for rotation about the axis and having a plurality of blades extending into the gas path. The turbine shroud segment includes a body extending axially between a leading edge and a trailing edge and circumferentially between a first and second lateral edges. The body has a radially outer surface and a radially inner surface. The radially outer surface includes a textured surface exposed to a cooling flow. The radially inner surface defines an outer flow boundary surface of the gas path next to a tip of one of the blades. A cooling flow passageway is defined in the body and extends axially between one or more cooling inlets receiving the cooling flow from the textured surface and one or more cooling outlets.

An assembly is provided for an aircraft propulsion system. This assembly includes a forward thrust duct and a thrust reverser system. The thrust reverser system includes a thrust reverser duct, a bullnose ramp and a plurality of protrusions. The bullnose ramp is adapted to provide a transition from the forward thrust duct to the thrust reverser duct when the thrust reverser system is in a deployed configuration. The protrusions are bonded or formed integral with the bullnose ramp. Each of the protrusions is adapted to interact with boundary layer fluid flowing along the bullnose ramp from the forward thrust duct into the thrust reverser duct when the thrust reverser system is in a deployed configuration.

An airfoil of a gas turbine engine includes a leading edge and an opposed trailing edge defining a chord between the leading edge and the trailing edge, wherein the chord has a chord length. A concave surface is between the leading edge and the trailing edge, which includes a first portion proximal the leading edge of the airfoil and a second portion proximal the trailing edge of the airfoil, wherein the first portion of the concave surface includes about 10% to about 50% of the chord length. An erosion-resistant ceramic, cermet or intermetallic coating is on the second portion of the concave surface, which includes a coating leading edge pattern. The first portion of the concave surface is free of the erosion-resistant coating.

An exhaust system for an engine includes an exhaust nozzle located adjacent an outlet end of the engine to receive a primary flow of exhaust gasses expelled from the engine, an inlet opening formed between the exhaust nozzle and the outlet end of the engine through which a secondary flow is provided to the exhaust nozzle, and a vortex generator arranged within the exhaust system at a position where both the primary flow and the secondary flow are present. The vortex generator interrupts at least one of the primary flow and the secondary flow.

Baffle inserts for airfoils of gas turbine engines are described. The baffle inserts include a baffle insert body having a first side portion and a second side portion, wherein each side portion has a respective end, a first set of vortex generation elements is arranged at the end of the first side portion, and a second set of vortex generation elements is arranged at the end of the second side portion. The first set of vortex generation elements and the second set of vortex generation elements are arranged at an aft end of the baffle insert body.

A cooling circuit for a gas turbine engine includes a gas turbine engine component having at least one internal cooling cavity defined by an internal wall surface and a plurality of turbulent flow features extending outwardly from the internal wall surface. Each turbulent flow feature is spaced apart from an adjacent turbulent flow feature in a first direction. At least one trench extends through the turbulent flow features in the first direction, and a plurality of cooling holes are formed within the at least one trench. A gas turbine engine and a method of forming a cooling circuit for a gas turbine engine component are also disclosed.